The present invention relates to an axial turbine for aeronautical applications and, in particular, for an aeronautical jet engine.
As is known, an aeronautical engine comprises a compressor unit, a combustion chamber arranged downstream from the compressor unit and a turbine unit, which is in turn arranged downstream from the combustion chamber and, generally, comprises three axial turbines, which are designated as high-, medium- and low-pressure turbines depending upon the pressure of the gas passing through them.
Each axial turbine comprises a succession of stages, each one of which consists of a stator comprising an array of fixed vanes and a rotor comprising an array of vanes that rotate about the axis of the turbine.
The efficiency of a known axial turbine and consequently of the associated aeronautical engine varies substantially as a function of the various operating conditions of the aeronautical engine itself.
Indeed, the flow rate and thus the velocity of the gas passing through the turbine stages vary as a function of engine operating conditions, while the geometry and relative position of the vanes of the stages themselves are set at the design stage in accordance with a fixed compromise configuration so as to achieve a satisfactory average efficiency for any gas flow rate and for any engine operating condition.
It has been found necessary to improve turbine efficiency and thus the overall efficiency of the associated aeronautical engine under the various operating conditions of the engine.